1. Field of the Invention
The present invention relates generally to the field of advanced composite aerostructure articles and methods of fabrication, more particularly, to an advanced composite aerostructure article having an integral co-cured fly away hollow mandrel and a method of fabrication.
2. Background Information
There is a growing trend in the aerospace industry to expand the use of advanced composite materials for a diverse array of structural and dynamic aerostructural applications because of the strength-to-weight advantage provided by composite materials. One particular application for the use of such advanced composite materials lies in the fabrication of advanced composite articles such as panels for nacelles for aircraft jet engine propulsion systems and for control surface components, such as spoilers. Such structural articles generally comprise inner and outer composite skins, which are formed from composite materials such as graphite or an aromatic polyamide fiber of high tensile strength that are embedded in a resinous matrix, e.g., epoxy, having a honeycomb core material interposed therebetween. One or more stiffening members may be affixed to the outer skin and covered with an inner skin for efficiently transmitting and/or reacting axial and/or bending loads to which the component is subjected. Aerostructure articles, such as spoiler components, have more than one external surface, i.e., a lower and upper surface, having curvatures and often taper, at an aft location, to a sharp edge. Manufacture of these articles present additional challenges to current techniques for fabricating composite structures.
There are two techniques currently employed for bonding through autoclave processing a composite stiffening member in combination with a composite structural panel: (1) the co-cured bonding method; and (2) the secondary bonding method. Both methods are disadvantageous in requiring costly non-recurring tooling and/or costly recurring manufacturing steps.
A typical composite sandwich panel intended for use as an aerostructure article is normally fabricated using two autoclave cured inner and outer composite skins that are formed by using a curing cycle with heat, pressure, and a unique tool for each skin. A sandwich panel is then made up using a composite bond jig, tool or fixture with the pre-cured face skin laid up on the bond jig tool followed by a ply of film adhesive, a honeycomb aluminum or non-metallic core of a given thickness, another ply of film adhesive and finally the previously pre-cured inner skin. The bond jig that is used to fabricate the sandwich panel is generally the same tool that was used to create the outer composite skin. A plurality of closure plies of uncured composite material are layed up and the assembled sandwich panel are cured during their final assembly stage. This sandwich panel is then vacuum bagged to the composite bond jig and again cured in an autoclave under high pressure and heat.
Thus, at least three very expensive and man-hour consuming cure cycles have gone into the fabrication of this exceptionally strong but lightweight composite/honeycomb core sandwich panel. At least two different and expensive tools are needed in this process. Manufacturing flow time is very long, energy use is high and the manufacturing floor space required is excessive.
The co-curing method envisions curing the composite inner and outer skins that are laid up with a layer of adhesive film and honeycomb core in one cure cycle in the autoclave. A co-cured panel is desirable in that it is less expensive to fabricate since only one bond jig tool is required, only one cure cycle is needed, it is less labor intensive, it requires less floor space to accomplish, and a much shorter manufacturing flow time is achieved. However, co-curing an aerostructure panel has never achieved widespread acceptance because of the large loss of panel strength and integrity that is lost due to the lack of compaction of the composite plies placed over and under the honeycomb core details. The composite plies dimple into the center of each core cell with nothing but the cell walls to compact the composite skins. The only way to overcome this knockdown factor is to add extra plies which creates both unwanted weight and excess cost. Thus, because of these constraints co-cured aerostructure panels are not widely manufactured in the aerospace industry.
There are other particular problems when a honeycomb core element is used to provide a stiffening element for an aerospace article such as a fan cowl. As Hartz et al described in U.S. Pat. No. 5,604,010 concerning a “Composite Honeycomb Sandwich Structure,” with a high-flow resin system, large amounts of resin can flow into the core during the autoclave processing cycle. Such flow robs resin from the laminate, introduces a weight penalty in the panel to achieve the desired performance, and forces over-design of the laminate plies to account for the flow losses. To achieve the designed performance and the corresponding laminate thickness, additional plies are necessary with resulting cost and weight penalties. Because the weight penalty is severe in terms of the impact on vehicle performance and costly in modern aircraft and because the flow is a relatively unpredictable and uncontrolled process, aerospace design and manufacture dictates that flow into the core be eliminated or significantly reduced. In addition to the weight penalty from resin flow to the core, it was discovered that microcracking that originated in the migrated resin could propagate to the bond line and degrade mechanical performance. Such microcracking potential has a catastrophic threat to the integrity of the panel and dictates that flow be eliminated or, at least, controlled.
Unfortunately, the use of honeycomb core as a stiffener for elements in an aerostructure article such as a fan cowl, or in a structural panel has other deleterious effects, two of the greatest drawbacks to aluminum core being its inherent significant cost and corrosion. To minimize galvanic corrosion of the core caused by contact with the face skins, isolating sheets are interposed between such aluminum core and the face skins. Also, the aluminum core has an inherent cost and also must be machined to a desired shape in an expensive process. The honeycomb core may also be subject to crush during manufacture and thereby limits the pressures that may be used in autoclave processing. Also, the honeycomb core if damaged in use has a spring back effect which makes the detection of such damage more difficult.
Another method for reinforcing mandrels for stiffener elements, such as hat sections, for aerospace advanced composite structural panels involves the use of a composite stiffening member formed over a polyimide foam mandrel which is fabricated by machining a core mandrel to a desired shape. Obviously, the machining of the core mandrel is expensive and time consuming and further introduces the problem of properly bonding the core mandrel to inner and outer skins.
A novel and useful technique for fabricating an advanced composite aerostructure article is described in U.S. Pat. No. 6,458,309, and co-pending U.S. application Ser. No. 10/142,490, filed May 5, 2002, both of which are incorporated by reference in their entirety herein.
A need has arisen for a practical method of readily producing stiffened, fiber-reinforced composite structures for aerospace applications which are cost and labor efficient and which save time in the fabrication process.